Multiple stage gas compressors, for example those used in aircraft gas turbine engines, are frequently operated over a wide range of conditions, such as varying load, rotational speed, input and output pressure, mass flow, etc., which must be accommodated in order to insure efficient and stable operation.
To those skilled in the art of compressor design, the so-called compressor design point is in fact the result of a series of structural, operational, and other design compromises necessary to allow the compressor to function within the expected operating environment. Aircraft gas turbine engines typically present some of the harshest design compromises, with the engine being required to achieve startup, ground idle, flight idle, full takeoff thrust, cruise thrust and a variety of intermediate power output levels as required during aircraft operation. Ambient air pressure and temperature also change as the aircraft climbs from sea level to cruise altitude, and the consequences of even a momentary compressor stall or other operating instability are highly undesirable.
Modern turbofan gas turbine engines typically utilize a two spool compressor configuration wherein a low pressure compressor having a plurality of individual rotor stages partially boosts the incoming air pressure, with a second, multi-stage high compressor providing the remainder of the pressure increase necessary. Such dual compressors, although turning on separate shafts, are closely coupled not only by the fact that the high compressor receives the output compressed air from the upstream low compressor, but also due to the fact that the outlet from the turbine section driving the high compressor shaft is received by a downstream low turbine section which turns the low compressor shaft.
Prior art engines of such two spool design have been built which can function under the range of operating conditions discussed above, although it has been found advantageous under certain conditions to divert or bypass a portion of the compressed air from the compressor section and route it overboard in order to prevent a flow mismatch between the low and high compressors which could induce an undesirable compressor stall condition. Emphasis on high performance and low fuel consumption have also given rise to the variable geometry compressor wherein the plurality of stator stages disposed intermediate the rotor stages of a multi stage compressor are adapted to be repositioned by an external actuation mechanism during operation of the compressor.
The variable stator states in the high or low pressure compressor typically involve an external unison ring and crank linkage secured to each individual vane which causes all the vanes in a single vane stage to be rotated simultaneously about the radial axis, thereby varying the angle of attack of each vane as well as the flow area of the nozzle formed between adjacent vanes for the compressed air flowing therethrough. By judiciously varying the position of the stator vanes and hence the angle of attack and nozzle flow area of each vane stage, it is possible to achieve more efficient compressor operation over a wider range of overall volume flow and without sacrificing a margin of safety between engine operation and the occurrence of a stall condition.
The control and positioning of such variable stator vanes is thus of prime importance in achieving reliable and efficient engine operation. Current prior art control methods schedule the desired position of individual stator vane stages as a function of the rotational speed of the corresponding compressor rotor. These positions are commonly determined at the corresponding steady state operating point and hence require some modification during transient load changes or other varying operating conditions. In general, closing of the stator vane nozzle area increases the safety margin of the overall compressor relative to the occurrence of a stall condition, but simultaneously decreases the overall compressor operating efficiency and airflow rate. Prior art systems, attempting to maintain both a sufficient stall safety margin and compressor efficiency, have required complicated modifying factors to be applied to the steady state vane positioning signal as the engine experiences operating transients.
One effect of these complicating factors on aircraft gas turbine engine operation is to slow the response time to a snap change of throttle position, for example acceleration from flight idle to takeoff thrust or vice versa. In order to ensure that sufficient compressor stall margin is not lost during such extreme transients, prior art fuel controllers either slow the engine and control response time to allow the stator vane controls to more closely track engine performance or displace the operating line downward sacrificing steady state performance for a larger transient margin.
In addition to this undesirable slowing of engine response time, the prior art systems are also subject to reduced functionality due to wear or damage to the linkages positioning the individual vanes. Since the prior art systems are open loop controllers driven by schedules, wear or mechanical backlash in the linkages opening and closing the individual vanes can result in a misalignment of .+-.2.degree. or more from the desired vane angle of attack. Such misalignment may be global, caused by inaccuracy of the positioning actuator, as well as local caused by backlash in the individual vane positioning cranks, etc. Such inaccuracies may, of course, increase as the linkage is worn and may even force the compressor closer to a stall condition under certain circumstances. Steady state performance must be degraded at design to allow for this anticipated deterioration.
What is required is a vane positioning method and system which monitors compressor performance by a feedback control loop to determine vane position error and to modify the vane positioning signal as necessary to achieve proper collective vane alignment.